Directional control-automatic meteorological compensation (d.c.-automet) inertial guidance system for artillery missiles

ABSTRACT

A missile system disposed for flight in a planned trajectory ending at a target. The trajectory includes booster propulsion and sustainer propulsion phases of flight. The missile is controlled in the booster propulsion phase by means of a directionally controlled (D.C.) inertial guidance system which controls thrust direction. Booster control means is provided for creating pressure differentials across the propulsive gases of the booster motor for thrust vectoring the missile to maintain the missile in the trajectory during the booster propulsion phase of flight. In the sustainer phase of flight, the missile is controlled by an automatic meteorological compensation inertial guidance system (AUTOMET) which controls thrust magnitude. Sustainer control means is provided for varying the thrust of the sustainer to maintain the difference in the actual and desired velocities at null and thus maintain zero net axial force on the missile in the sustainer phase of flight. The system is, therefore, known as the D.C.-AUTOMET guidance system. Previous inertial guidance systems utilized a stabilized platform having a number of accelerometers mounted thereon. A plurality of gyroscopes were needed to provide a reference for the stabilized platform. Furthermore, in guidance systems of this type, very expensive precision instruments were required to perform the guidance and control functions.

United States Patent [72] Inventor William C. McCorkle, Jr.

Huntsville, Ala. [21] Appl. No. 540,480 [22] Filed Apr. 4, 1966 [45]Patented Mar. 9, 1971 [73] Assignee The United States of America asrepresented by the Secretary of the Army [54] DIRECTIONALCONTROL-AUTOMATIC METEOROLOGICAL COMPENSATION (D.C.- AUTOMET) INERTIALGUIDANCE SYSTEM FOR ARTILLERY MISSILES 9 Claims, 5 Drawing Figs.

[52] US. Cl 244/3.2, 244/3.22 [51] Int. Cl F42b 15/18 [50] FieldofSearch 244/3.14, 3.15, 3.19, 3.20, 3.22, 3.23, 3.21

[56] References Cited UNITED STATES PATENTS 2,979,284 4/1961 Genden eta1 244/3.14 3,008,668 11/1961 Darlington 244/3.11 3,184,182 5/1965 Mayet a1. 244/3.22 3,249,325 5/1966 Forehand 244/3.22

Primary Examiner-Verlin R. Pendegrass Attorneys-Harry M. Saragovitz,Edward J. Kelly, Herbert Earl and Harold W. Hilton ABSTRACT: A missilesystem disposed for flight in a planned trajectory ending at a target.The trajectory includes booster propulsion and sustainer propulsionphases of flight. The missile is controlled in the booster propulsionphase by means of a directionally controlled (D.C.) inertial guidancesystem which controls thrust direction. Booster control means isprovided for creating pressure differentials across the propulsive gasesof the booster motor for thrust vectoring the missile to maintain themissile in the trajectory during the booster propulsion phase of flight.In the sustainer phase of flight, the missile is controlled by anautomatic meteorological compensation inertial guidance system (AUTOMET)which controls thrust magnitude. Sustainer control means is provided forvarying the thrust of the sustainer to maintain the difference in theactual and desired velocities at null and thus maintain zero net axialforce on the missile in the sustainer phase of flight. The system is,therefore, known as the D.C.-AUTOMET guidance system.

Previous inertial guidance systems utilized a stabilized platform havinga number of accelerometers mounted thereon. A plurality of gyroscopeswere needed to provide a reference for the stabilized platform.Furthermore, in guidance systems of this type, very expensive precisioninstruments were required to perform the guidance and control functions.

PATENTEDW em 3568.954

SHEET 2 BF 2 48 w 44 THRUST VECTOR VALVE GYR E R i 00 i VALVE t VALVESIGNAL AMPLIFIER DRIVER SELECTOR THRUST VECTOR VALVE 48 FIG. 4

l8 2 sa: ACCELEROMETER INTEGRATOR I CONTROL VALVE 26 27 56 PRESETBOOSTER COMPARATOR CUTOFF VELOCITY VALVE FIG. 5

FLIGHT PATH ALTITUDE William C. McCorkle,Jr.,

INVENTOR. Ll DOWNRANGE M m. M

s DISPLACEMENT BY M W FIG. 3 )m'dm DHRECTHONAL CONTROLAUTOMATICMETEORULGGHCAL C(lMlPlENSATlON (lD.C.-AUTOMET) lll lE ilTllAL GlJlDANClE SYSTEM FOR ARTILLIERY MlSlllLES The present invention overcomes theabove noted defects by providing an inertial guidance system in whichthe requirement for a stabilized platform is removed and the requirementfor gyroscopes to provide a reference for the stabilized meterologicalis also removed.

In describing the present invention, some principle sources ofdispersion in tin stabilized missile will be explained first. A typicalunguided fin stabilized rocket is launched form a short rail or tube, isaccelerated to the required velocity by the reaction of the rocket jetduring the burning period, and afterward coasts to the target in thesame manner as other fin stabilized ballistic projectiles. Dispersionoccurs because of a number of factors.

For example, during the launching operation, a small angular motionabout a transverse axis is imparted to the rocket; part of this angularmotion is repeatable and causes no error if allowance is made, but partis random and causes dispersion. This error source is called mallaunch.

Also, after launch the thrust may fail to pass precisely through therocket center of gravity and a moment tending to rotate the rocket abouta transverse axis is thus produced. As a rotation from this cause (orany other, such as mallaunch) occurs, the rocltet attitude deviates fromthe desired direction and the thrust drives the rocket off the intendedcourse. This error source is called thrust malalignment and is largely,but not entirely, due to manufacturing tolerances or imperfections. Alsoduring the burning phase the rocket responds to winds which have acomponent normal to the flight path, called cross wind. Since the rocketis stable, it tends to rotate into the wind, and the thrust drives itoff the intended path in the upwind direction. This effect ispredictable, but proper allowance for it calls for measurement of thewind. The difference between the wind as assessed by measurement priorto flight and that actually experienced by the rocket during burningfurther contributes to dispersion.

Furthermore, after burning, the rocket behaves as a tin stabilizedprojectile. Dispersion is caused by the variations in the velocityachieved, and to a much lesser extent, by variations in distancetraveled during burning. The velocity error is exactly analogous to thevariations in muzzle velocity in a cannon, with the same result rangedispersion. The errors in direction of the flight path accrued duringburning are propagated through the coast phase and contribute both torange and azimuth dispersion. Also, variations in weight or dragcoefficient of the rocket after burning cause range dispersion. Mostimportant, atmospheric conditions along the path of the rocket to thetarget affect the path and must be taken into account. Winds andnonstandard atmospheric density perturb the normal trajectory, and theseperturbations can be predicted if the winds and density variations areknown. To the extent that the assessment of these meteorological factorsis different from that actually experienced by the projective in itsflight, the compensation will be incorrect and further range and azimuthdispersion will accrue.

The rocltet may have bent or misaligned fins which can cause dispersion,such dispersion arises during burning, and may increase after burnout.However, a slow roll is generally imparted during the burning period toreduce the effect of thrust malalignment, and the rolling motionpersists (in practice it is usually maintained by a slight cant to thefins) during the coast phase. A roll of sufficient rate to appreciablyreduce the effect of thrust malalignment is more than enough to make theeffect of any aerodynamic asymmetry completely negligible (provided theroll rate and natural yawing frequency do not coincide or lie too closeto each other).

in summary, the principal sources of dispersion in fin-stabilizedrockets during burning are mallaunch, thrust malalignmerit, andmalassessrnent of cross wind; after burning, variations in the velocityimparted during burning, variations in drag coefficient or burnout mass,and malassessment of winds and density along the trajectory. There areother lesser sources of error which will be mentioned in describing theinvention claimed. I

It is, therefore, an object of this invention to provide the minimumguidance to the free rocket necessary to make it insensitive to thedisturbances discussed above, and to remove the necessity for makingmeasurements of the meteorological factors.

It is another object of the present invention to provide a missilesystem having a guidance system composed of a few relatively simple andinexpensive components and having a high degree of accuracy.

It is a further object of the present invention to provide such aguidance system to control thrust direction during the boosterpropulsion phase of flight and thrust magnitude during the sustainerphase of flight.

A still further object of the present. invention is to provide such aguidance system which does not require the use of a stabilized platformand reference gyroscopes therefor.

A yet further object of the present invention is to provide a guidancesystem in which a single accelerometer is utilized for terminatingbooster thrust, initiating sustainer phase of flight, and controllingsustainer thrust magnitude to maintain the missile in substantially avacuum trajectory.

To accomplish these and other objects, the present invention includes amissile disposed for flight along a predetermined trajectory to atarget. The trajectory includes booster propulsion and sustainerpropulsion phases of flight. To maintain the missile in the desiredtrajectory at the proper velocity during the booster phase, the missileis provided with a directional control inertial guidance system. Thisdirectional control inertial guidance system is comprised of a gyrodisposed for generation of an electrical signal indicating pitching andyawing of the missile. The signals actuate a plurality of injectionvalves disposed around the periphery of the motor for injecting fuelinto the nozzle, by secondary injection, to produce moments which tendto restore the missile axis to the gyro spin axis.

The sustainer phase of flight is controlled by an accelerometer whoseinput axis is oriented along the longitudinal axis of the missile. Theelectrical signal from the accelerometer is integrated to obtain theelectrical analogue of missile velocity. This velocity signal terminatesthe boost phase and initiates the sustainer phase upon reaching thedesired preset velocity corresponding to the desired range. Theaccelerometer also provides electrical signals which are utilized tovary the magnitude of thrust of the missile during the sustainer phaseof flight. The thrust is varied in response to axial accelerationsacting on the missile to maintain zero net axial force on the missileand thus cause the missile to fly in a vacuum trajectory.

The above mentioned advantages and objects of my invention will becomemore readily apparent from the following detailed description, takentogether with the accompanying drawings in which:

FIG. 1 is a partially cut away elevational view of a missile embodyingmy invention.

FIG. 2 is a sectional view along line 2-2 of FIG. 1 illustrating theconcentric arrangement of the sustainer and booster motors of thepresent invention.

FIG. 3 is a diagrammatic view of a missile and the desired flighttherefor and illustrating the coordinate system.

FIG. 4 is a schematic diagram showing the relationship of the gyro andthrust vector valves for the booster phase of flight.

FIG. 5 is a schematic diagram showing the relationship of theaccelerometer and throttling valves for the sustainer phase of flight.

in describing the invention, a complete firing operation will befollowed through from the laying operation to impact.

It is assumed target range and azimuth information are available, andthat certain corrections due to earths rotation, launcher and targetrelative altitude, etc. have been provided.

A missile is elevated to a predetermined angle for travel in atrajectory. The angle may be the same for all firings. For maximumrange, this angle is 50 to 55. The angle may be determined externally orinternally by a preset level indicating device, many variations of whichare known to the art.

As is shown in FIG. I, the missile includes a gyro 12 which is mountedin the missile axis and is aimed both in azimuth and elevation. The mostaccurate method of performing this function would involve a reflectingsurface on the gyro rotor itself, oriented so as to be normal to thetrue spin axis. Numerous alternate approaches exist, all with varyingpotential accuracies. In any event, by standard artillery procedures,the gyro spin axis is turned to the proper azimuth, and elevated to therequired elevation, taken to be 50 for descriptive purposes here.

For control of range, a main boost motor 14 is provided and thrust ofmotor 14 is terminated at a predetermined speed, which is, except forminor correction,

where R is range, 3 is earths gravity and 9 is the boost flight pathangle measured from the horizontal.

The preset velocity corresponding to the desired range is set into themissile guidance system, as a number for digital circuits, or as avoltage for electrical anolog techniques. A procedure for performingthis operation is as follows: with missile 10 on a launcher (not shown)at firing elevation, an accelerometer 18, is set with its sensitive axisprecisely oriented along the missile axis 20, or alternatively, ismounted so that its sensitive axis may be turned to the vertical. Theobject is for accelerometer 18 to read a precisely known fraction ofearth 3, as determined by the inclination of the sensitive axis from thevertical. By either knowing the angle precisely or turning it to thevertical (less precision is required for the latter), the precise inputto the accelerometer, in terms of earth g, is known. The accelerometeris connected to an integrator 22 and the accelerometer and integratorare activated, and the accelerometer output is integrated for aprecisely known time, say 10 seconds. The output of integrator 22 as anumber or voltage, is read out at the umbilical connection to themissile 24. The output represents 10 g sin 9 second, where 6 is theangle of elevation of the accelerometer sensitive axis from thehorizontal. For 6 of 50, 10 g sin 9 represents a velocity of 75meters/second. Suppose the desired range is 25,000 meters, whichrequires a boost velocity of 495 meters/second. Then the output ofumbilical connection 24 of the 10 g sin 6 input, which represents 75meters/second, must be multiplied by 6.6 to obtain a magnituderepresenting 495 meters/second. This is done, and the new number orvoltage is inserted back into the missile via umbilical connector 24 toa velocity cutoff, or range presetting device 26.

The missile is now ready to be fired. The azimuth, elevation, and rangesettings are completed. The firing sequence is initiated by an input viaumbilical connector 24 which activates the missile power supply 28,spins up the rotor of gyro 12 and ignites a gas generator 30. The gasfrom the gas generator pressurizes a pair of pistons 32 and 34 in liquidfuel and oxidizer tanks 36 and 38, and starts the flow of propellantsthrough appropriate frangible seals and plumbing to the booster andsustainer motors l4 and 40, respectively. Motors 14 and 40 areconcentrically disposed in the missile. The propellants are hypergolic(for example, inhibited red fuming nitric acid and unsymmetricaldimethyl hydrazine) which ignite in the two chambers, and thrust isproduced. The gas pressure on the pistons 32 and 34 is regulated by apressure regulator 42, and the excess pressurizing gas is dischargedoverboard through two diametrically disposed tubes 43. The first motionproduced by the resulting thrust closes an acceleration switch inintegrator 22, and the output of accelerometer l8 begins to beintegrated to produce a voltage (or number) proportional to velocity,minus the component of integrated earth gravity along the missile boostflight path. The ground connection to the missile guidance system viaumbilical connector 24 is released just prior to first motion, and nofurther input to the missile is made. After a short travel on thelauncher rails, the missile is in flight on its own.

Steering of the missile during the boost phase is accomplished by gyrol2 and appropriate steering forces as shown in FIG. 4. The signals fromthe gyro include two signal outputs to two control channels indicatingpitch and yaw of the missile axis relative to the gyro. Each channel mayinclude in an electronic package 44, (FIGS. 1 and 4), a DC signalamplifier, valve driver, valve selection circuit and two injectionvalves 48. The gyro error signal is amplified by the amplifier and thenpasses through the valve driver and selection circuit to the propervalve through conduits 46. Valve selection depends upon the algebraicsign of the error signal. Valves 48 inject fuel into the nozzle in sucha manner as to produce deviations in the rocket thrust in proportion tothe gyro signals in each body-fixed plane, initially designated pitchand yaw, but rolling with the missile as it rolls. The secondaryinjection, by techniques well known in the art, produce moments whichtend to restore the missile axis to the gyro spin axis. If the thrustdeviation angle is denoted by 6, and the angle between the gyro axis(reference direction) and missile axis is denoted by p, (FIG. 3) thenthe ratio 8/p is called 'y, the gain. This relation exists in both thepitch and yaw body fixed planes, and the value of 'y is chosen so as tominimize the flight path deviation of the missile during the boostphase. It can be shown that the minimum response of the missile flightpath to cross wind disturbances occurs when 7 is chosen so as toneutralize the upwind and downwind tendencies of the missile. If y ischosen too large, the attitude of the missile is maintained too rigidly,and a downwind drift results. If, on the other hand, 7 is chosen toosmall, the missile tends to weathercock into the wind too much, and thethrust drives the missile upwind. If the steering force is located atthe missile nozzle, it can be shown that the optimum value of 'y isgiven by where 1,, is the moment arm of the aerodynamic center ofpressure about the center of mass of the missile, and 1 is the momentarm of the control forces about the center of mass. It must be pointedout that a stable missile, i.e., one with the aerodynamic center ofpressure aft of the center of mass, is necessary for proper functioningof this guidance system.

Because the quantities 1,, and 1 vary during the boost phase, due to themovement of the center of mass as propellant is consumed, and the shiftin aerodynamic center of pressure as the Mach number increases, theoptimum value of y changes during boost. If the accuracy requirementspermit, a value of representing a best average may be chosen;alternatively 7 may, by suitable programming of the electricalparameters, be made to follow the optimum value.

The choice of y is dictated by the requirement for removing thesensitivity to cross wind present in the free rocket. The proper choiceof y results in virtually complete insensitivity to cross winds, whethersteady or changing in distance and time. But there are otherdisturbances, and the optimum choice of 'y for cross winds in notoptimum for the other prinicipal source of free rocket disturbance,thrust malalignment.

The optimum value of y for reducing thrust malalignment caused by an offaxis CG. is infinity-the stiffest possible control in attitude. Ratherthan compromise the wind sensitivity, however, the gain, 7, is kept atthe optimum value for minimum wind sensitivity, and the effect of thrustmalalignment is further controlled, beyond that achieved by the crosswind optimized yby rolling the missile. The flight path deviation withno roll, for unit thrust malalignment angle, is 1/7. Since typicalvalues of y are .3 to .4, this sensitivity, while much smaller than thatfor free rockets, is still unsatisfactorily large. If the rocket rollsthrough an angle l1 during the boost phase, however, the sensitivity tounit thrust malalignment is reduced to l/yul, provided the roll rate isreasonably uniform, and several complete revolutions are turned.Typically at least live or more complete revolutions will be turnedduring burning, and ab has a value of 30 or more radians. The roll rateneed not be uniform to be effective, but the magnitude of the reductionin thrust malalignment effect depends on the particular roll ratehistory which occurs during burning.

The greatest effect of roll rate is achieved if the desired uniform rollrate is present at the time of release of launcher constraints. This maynot be practicable, and an alternativeis a slow buildup of roll ratecoupled .with a change in the valve of 'y from ahigh initial value tothe value appropriate for minimum wind sensitivity after the roll rateis established, but before aerodynamic forces become very large. Thetime of change in -y should be before a flight speed of .4 or .5 Machnumbers is reached. This approach takes advantage of the fact that atlaunch aerodynamic forces, and hence cross wind effects, are at aminimum, while the effects of missilemisalignments are maximum at thistime. Thus thevalue of y can be raised to reduce the effects of'misalignments' at this time, and lowered to compensate for wind effectswhen they become important as speed builds up. In the interval, thebuildup of roll rate prevents any substantial effect of misalignmentsafter y is reduced.

One way of producing the desired roll" is to put fixed jet vanes 50 inthe nozzle. expansion cone, set at an angle, to produce a roll torque.This torque would result in a continuous roll acceleration, except thatas the missile velocity increases the stabilizing fins 52 ten to opposeroll rates greater or less than some rate proportional to the missileforward velocity and depending on their cantangle. The resultant of thejet vane torquesand stabilizing fin torques tend to produce a roll, raterising rapidly at first, then perhaps decreasing slightly as the finstake effect, and finally increasing again in proportion to forwardvelocity as the fin cant torques become dominant.

, Other rolling techniques may be used. For example, small auxiliaryspin rockets, firing tangentially to produce a roll torque immediatelyafter launch, represent an acceptable technique well known to the artand widely utilized.

Because of the rolling motion, it is important that the controlrestoring moments follow closely in phase and amplitude the gyrosignals, lest an error detected in one channel provide,

because of the rolling, a corrective force in the other channel.

This difficulty will not appear if there is no phase lead or lag ofoutput control forces in relation to control system inputs (i.e.

gyro pitch and yaw signals) at the frequencies corresponding to missileroll rates. This requirement is apart from the phase and amplituderelations necessary for stabilization of the control system, affectingeither the stability of subsystems or the missile pitch and yawoscillations caused by control action. It is worth noting that, becauseof the relatively short duration of the boost phase, and the steadilyincreasing aerodynamicstability with missile velocity, control systemstability in the classical sense of the location in the complex plane ofthe roots of the characteristic equation governing missile pitch andyaw, evaluated for fixed values of the coefficientsin the equation, maynot be relevant. in fact, entirely satisfactory operation of the boostphase directional guidance system may be obtained even though the systemis unstable in the sense described axis is the intended direction ofaim.

During the boost phase, gravity acts in the pitch plane to produce adownward bias. The missile control guidance system functions to causeacceleration to occur only along the direction of the gyro spin axis,regardless of external or internal disturbances. Superimposed on thisacceleration along the gyro axis is, however, the gravitationalacceleration in the vertical direction. The flight path acceleration isthe resultant of missile acceleration along the gyro axis and thevertical acceleration due to gravity, added vectorially.

Because the missile tends to hold an attitude parallel to the gryo axis,the velocity perpendicular to the gyro axis resulting from the effect ofgravity appears to the missile as an upward wind of magnitude gtsin6,'where 9 is the elevation of the gryo axis. This steadily increasinggravity induced wind causes insignificant deviation from the flight pathdetermined by the resultant of missile and gravitational acceleration,since the control gain y, as previously described, has been optimized tominimize the flight path deviation due to cross wind.

Because of the effect of gravity, it is important that the missile axialacceleration, or axial force to missile mass ratio be reproducible frommissile to missile, since the magnitude of the downward bias due togravity is proportional to the ratio of missile acceleration togravitational acceleration. If this ratio is not sufficientlyreproduciblefor the accuracy requirements, further steps may be takenwhich will be described in detail below. l

The directionally controlled boost phase continues until the output ofaccelerometer 18, after integration by integrator 22 to obtain missilevelocity, becomes equal' to the preset velocity, corresponding to thedesired range, stored as a number or voltage at velocity cutoff device26. The output of integrator 22 is compared with the preset velocity ina com parator 27 and the booster cutoff signal is initiated at, orperhaps slightly before the preset velocity value to allow for lags andthrust decay, the'time the missile velocity reaches the preset velocity,and is transmitted via conduit 54 to a cutoff valve 56, which stopspropellent flow to the boost motor and secondary injection valves 48. i

The directional control boost phase is terminated at this point, and thesustain phase is initiated. Up until this time, the sustainer motor 40has been operating at full thrust.

The sustain phase utilizes the Automatic Meteorological Compensation(AUTOMET) technique. As shown in FIGS. 1

and 5 an output of integrator 22 which essentially represents thedifference between the preset velocity at 26 and the missile velocityfrom the accelerometer l8 and intergrator 22, is used to operatea'sustainer control valve 58 connected intermediate motor 40 and thefuel and oxidizer tanks. Control valve 58 is a throttling valve. If themissile velocity is higher than the preset velocity, the valve closes ata rate proportional to the difference. On the other hand, if the missilevelocity is lower than the preset velocity, the valve opens at a rateproportional to the difference. In order to stabilize this controlsystem, the signal representing the difference velocity from integrator22 mayalso contain terms proportional to the axial acceleration, and ofthe proper sign to provide sustainer control system damping,

During the sustain phase, which occupies most of the total flight timeof the missile, the difference between the preset velocity and missilevelocity and missile velocity as representedby the integratedaccelerometer output, is kept at null by the sustainer thrust. Thismeans that the integral of the axial acceleration is kept constant, andthat the average net axial force is zero. Because the average net axialforce is zero during the sustain phase, or any appreciable part of it,the missile flight path is substantially the same as if it were in avacuum. There will be transient excursions of the net axial force fromthe balance of thrust and drag, but these produce negligible effect aslong as the net axial force, averaged over a small fraction of the totalflight time, is zero. t

It is clear why variations of atmospheric density or missile drag, orwind components along the flight path, cause no i change in theflightpath. These effects show up directly as increased or reducedthrust levels to keep the integrated accelerometer output at-the presetvalue. For components of 7 wind normal to the intended vacuum flightpath, however, the reason why the flight path remains essentiallyunaltered is less obvious. Consider the direction of air flow over themissile body after encountering a cross wind. It is composed of a largecomponent along the flight path direction because of the missilevelocity, and a component normal to the flight path due to the crosswind. Because the missile is aerodynamically stable, it tends to alignitself with the resultant direction of the air flow due to cross windand motion along the flight path. In other words, only transient lateralforces act on the missile during the time of alignment with the airflow, and the motion is of a damped oscillatory nature whose integratedeffect is very small, After alignment with the air flow caused by thecross wind and forward velocity,-there exists no steady lateral force onthe missile, and the control system holds the average net axial force atzero by varying thrust. Thus the missile center of mass continuesundisturbed along the vacuum flight path, while the missile attitude isslightly offset from the tangent to the flight path to maintainalignment with the air flow. A common analogous situation is that of anaircraft holding an upwind heading to compensate for cross wind, or aboat head- .ing upstream in order to make a direct crossing of a stream.

In order minimize during the sustain phase the effect of aerodynamicasymmetries which might produce a body fixed steady lateral force, theroll initiated during the boost phase is continued throughout flight bythe cant built into the stabilizing fins. The roll rate is notimportant, except that it is desirable to keep it above the naturalyawing frequency of the missile to avoid amplification of the yawingmotion when roll and yawing frequencies lie too near each other. Thereason for choosing the roll rate above the yawing frequency is that thelatter may drop to very low values when the missile is traversingrarefied atmosphere, while the equilibrium roll rate is not affected byair density. Hence if the roll rate is initially less than the yawingfrequency, they may pass through each other during the flight.

lt should be pointed out that the actual missile velocity along theflight path is not constant, even though the integrated accelerometeroutput is kept constant by the thrust control system. This is becausethe accelerometer sensitive mass and all other parts of the missile areequally affected by gravity, hence the missile changes velocity due togravity with no indication appearing at the accelerometer output. Thehorizontal velocity component remains constant, but the verticalvelocity component after the boost phase is steadily reduced by gravity,passing through zero at the summit of the trajectory, then increasing inthe downward direction, in exactly the same manner as any ballistictrajectory in vacuo.

For ranges near the maximum range of the missile, as determined by thepropellant capacity and gross missile weight, the sustain phase of theflight may be terminated by timer 60. Because of the remainingrelatively short time to target impact, the missile will deviate littlefrom a predictable trajectory if the AUTOMET control is terminated, buta relatively large amount of propellant will be saved, therebypermitting a missile of less gross weight than if operation weremaintained to impact. The actual missile velocity in the terminal phaseapproaches the velocity at booster cutoff, and this, together with theincreasing air density at lower levels, requires steadily increasingsustainer thrust to overcome drag and maintain zero axial acceleration.The duration of the free flight terminal phase will depend on therequired accuracy, with a less stringent accuracy requirement permittingmore terminal free flight.

At ranges less than maximum range, the free flight phase will beshorter, and at all ranges less than a fraction of the maximum range,depending on the free flight duration chosen for maximum range, thesustain phase may last until impact.

it was mentioned above that it was assumed that the booster thrust levelwas sufficiently reproducible to permit satisfactory operation, from theaccuracy standpoint. This may not always be true, depending on theaccuracy requirements and the booster reproducibility. The requirementfor booster thrust history reproducibility stems from several separateeffects on accuracy. One concerns the effect of gravity mentionedpreviously, in that the bias in the vertical plane depends on the ratioof average boost acceleration to the component of gravity normal to theintended boost flight path. This effect will not result in significantimpact error if the boost flight path has an elevation near that whichproduces maximum range for a given boost velocity, i.e., 45. It will,however, affect the time of flight to impact.

The component of gravity along the flight path has the effect ofreducing the actual velocity from that indicated by the integratedaccelerometer output, sometimes called the inertial velocity, If theaverage boost acceleration varies from missile to missile, then the timeto teach the preset inertial velocity must vary, hence the actualmissile velocity will differ by varying amounts from the preset velocitybecause of the varying time of action of the gravity component. Thiswill be partly compensated by the fact that the actual distancetraversed during the boost phase will vary in inverse proportion to theaverage boost acceleration, but there will nevertheless be a net errordue to booster average thrust variation.

A method of compensation for lack of booster thrust reproducibility, orto improve reproducibility, takes advantage of the variablethrustsustainer and control system during the boost phase. A voltage or numberis generated by a velocity increase program mechanism 62, connected tointegrator 22. The generated voltage or number is proportional to andincreases with time approximately as the expected missile velocity underthe action of the nominal booster thrust plus one-half of the sustainermaximum thrust. The output of mechanism 62 starts'from zero and isinitiated by the same switch which starts the integrator 22 to integratethe output of accelerometer 18 with first motion of the missile. Theoutput of mechanism 62 is actually a function which is the time integralof the quantity thrust minus drag divided by the mass (which decreasesuniformly as propellant is consumed), and is easily approximated to therequired accuracy by several straight line segments. The differencebetween the signal from mechanism 62, which is a programmed velocityincreased during boost, and the measured velocity from accelerometer 18and integrator 22 is used to control the sustainer thrust in the samemanner as is described supra during the sustain phase. The basicdifference is that during the sustain phase, the inertial velocity(i.e., output of integrator 22) is controlled to the constant presetvelocity, while during the boost phase, it is controlled to a steadilyincreasing velocity program supplied by mechanism 62 which duplicatesthe nominally expected velocity program under the action of the boosterand one-half the sustainer thrust. When the preset velocity on cutoffdevice 26 is reached, the booster cutoff signal is also used todisconnect mechanism 62 and connect cutoff device 26 so that thesustainer thrust is subsequently regulated to keep the differencedifference between cutoff device 26 and the output of integrator 22 atnull. Clearly, this system permits compensation for high or low valuesof the booster thrust up to one-half the maximum thrust of thesustainer, which is assumed variable from off to full thrust.

It should be readily apparent that applicant has provided a missilesystem having the accuracy and range of a guided missile with the lowcost and reliability approaching that of a free rocket.

Obviously, numerous modifications and variations of the presentinvention are possible in light of the above teachings. It is,therefore, to be understood that within the scope of the appended claimsthe invention may be practiced, otherwise, than as specificallydescribed herein.

Iclaim:

l. A missile disposed for flight according to a planned trajectoryending in a target including booster propulsion and sustainer propulsionphases of flight, said missile provided with a booster motor disposedfor providing substantially constant thrust during operation thereof,and a sustainer motor disposed for providing thrust during the boosterand sustainer 9 phase of flight with the magnitude of the thrust beingvariable in the sustainer phase; a gyro mounted in the missile with thespin axis along the longitudinal axis of the missile, said gyro disposedfor providing a signal voltage proportionate to the amount of deviationof said missile axis from said spin axis responsive to said deviations;booster control means disposed for actuation by said gyro signal voltagefor creating pressure differentials across the propulsive gases of saidbooster motor for thrust vectoring of said missile to restore saidmissile axis in alignment with said gyro spin axis during said boosterpropulsion phase of flight; accelerometer means mounted along saidlongitudinal axis of said missile for producing a signal voltageproportionate to the difference between the actual missile velocity anda desired velocity, said accelerometer signal disposed for terminatingbooster thrust when said difference in actual and desired velocity isnull; sustainer control means actuatable by said accelerometer signalsfor varying the thrust of said sustainer to maintain the difference insaid actual and desired velocities at null and thus maintain zero netaxial force on said missile in the sustainer phase of flight.

2. A missile as in claim 1 wherein said booster control means includes aplurality of injector valves mounted around the periphery of saidbooster motor and connected to a source of fluid under pressure, saidvalves disposed for selective actuation to control the direction ofthrust from said booster, and actuating means disposed for actuation ofsaid valves in response to signals from said gyro.

3. A missile as in claim 2 wherein said actuating means includesmechanism electrically connected to said gyro to receive error signalstherefrom in pitch and yaw and to selectively actuate said injectorvalves for diverting the direction of thrust of said booster motor tocorrect for said errors.

4. A missile as in claim 3 wherein said booster motor of said missile isprovided with a plurality of fixed jet vanes disposed therein, said jetvanes mount at an angle to the longitudinal axis of said missile andprojecting interiorly of said booster motor to produce a roll torque onsaid missile; and, fins mounted on the outer periphery of said missileand canted at an angle to the longitudinal axis of said missile formaintaining a predetermined roll of said missile in response toinitiation of said roll torques.

5. A missile as in claim 4 with said missile including a pair ofpropellant tanks connected to said booster and sustainer motors, saidtanks provided with pistons disposed for movement therein for forcingsaid propellants. to said booster and sustainer motors; and, means formoving said pistons.

6. A missile as in claim 4 wherein said sustainer control means includesa throttling valve communicating with said sustainer motor and saidpropellants to proportionately control flow of said propellants to saidsustainer motor to vary the thrust thereof in response to the differencein thrust and drag of the missile; said throttling valve being disposedfor actuation by said accelerometer output signal voltage.

7. A missile as in claim 6 wherein said sustainer motor is disposed insaid missile with its longitudinal axis coincident with the longitudinalaxis of said missile.

8. A missile as in claim 7 wherein said booster motor is concentricallydisposed about said sustainer motor.

9. A missile as in claim 8 wherein said means for moving said pistonscomprises a gas generator disposed for ignition for producing gases forexerting pressure on said pistons for movement thereof to forcepropellant from said propellant tanks.

1. A missile disposed for flight according to a planned trajectoryending in a target including booster propulsion and sustainer propulsionphases of flight, said missile provided with a booster motor disposedfor providing substantially constant thrust during operation thereof,and a sustainer motor disposed for providing thrust during the boosterand sustainer phase of flight with the magnitude of the thrust beingvariable in the sustainer phase; a gyro mounted in the missile with thespin axis along the longitudinal axis of the missile, said gyro disposedfor providing a signal voltage proportionate to the amount of deviationof said missile axis from said spin axis responsive to said deviations;booster control means disposed for actuation by said gyro signal voltagefor creating pressure differentials across the propulsive gases of saidbooster motor for thrust vectoring of said missile to restore saidmissile axis in alignment with said gyro spin axis during said boosterpropulsion phase of flight; accelerometer means mounted along saidlongitudinal axis of said missile for producing a signal voltageproportionate to the difference between the actual missile velocity anda desired velocity, said accelerometer signal disposed for terminatingbooster thrust when said difference in actual and desired velocity isnull; sustainer control means actuatable by said accelerometer signalsfor varying the thrust of said sustainer to maintain the difference insaid actual and desired velocities at null and thus maintain zero netaxial force on said missile in the sustainer phase of flight.
 2. Amissile as in claim 1 wherein said booster control means includes aplurality of injector valves mounted around the periphery of saidbooster motor and connected to a source of fluid under pressure, saidvalves disposed for selective actuation to control the direction ofthrust from said booster, and actuating means disposed for actuation ofsaid valves in response to signals from said gyro.
 3. A missile as inclaim 2 wherein said actuating means includes mechanism electricallyconnected to said gyro to receive error signals therefrom in pitch andyaw and to selectively actuate said injector valves for diverting thedirection of thrust of said booster motor to correct for said errors. 4.A missile as in claim 3 wherein said booster motor of said missile isprovided with a plurality of fixed jet vanes disposed therein, said jetvanes mount at an angle to the longitudinal axis of said missile andprojecting interiorly of said booster motor to produce a roll torque onsaid missile; and, fins mounted on the outer periphery of said missileand canted at an angle to the longitudinal axis of said missile formaintaining a predetermined roll of said missile in response toinitiation of said roll torques.
 5. A missile as in claim 4 with saidmissile including a pair of propellant tanks connected to said boosterand sustainer motors, said tanks provided with pistons disposed formovement therein for forcing said propellants to said booster andsustainer motors; and, means for moving said pistons.
 6. A missile as inclaim 4 wherein said sustainer control means includes a throttling valvecommunicating with said sustainer motor and said propellants toproportionately control flow of said propellants to said sustainer motorto vary the thrust thereof in response to the difference in thrust anddrag of the missile; said throttling valve being disposed for actuationby said accelerometer output signal voltage.
 7. A missile as in claim 6wherein said sustainer motor is disposed in said missile with itslongitudinal axis coincident with the longitudinal axis of said missile.8. A missile as in claim 7 wherein said booster motor is concentricallydisposed about said sustainer motor.
 9. A missile as in claim 8 whereinsaid means for moving said pistons comprises a gas generator disposedfor ignition for produCing gases for exerting pressure on said pistonsfor movement thereof to force propellant from said propellant tanks.